The present invention relates generally to gas turbine engines, and, more specifically, to turbine blade cooling therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the gases in a high pressure turbine (HPT) which powers the compressor, and in a low pressure turbine (LPT) which typically powers a fan in an aircraft turbofan engine application.
The HPT includes a stator nozzle disposed at the outlet end of the combustor which first receives the hottest combustion gases therefrom. The nozzle includes a row of vanes which direct the combustion gases into a row of first stage turbine rotor blades mounted to the perimeter of a supporting rotor disk.
The stator vanes and rotor blades have specifically configured airfoil profiles with generally concave pressure sides and generally convex opposite suction sides extending in radial span, and extending in chord between opposite leading and trailing edges. The airfoils are hollow and include various cooling circuits therein for channeling therethrough a portion of pressurized air bled from the compressor for cooling the airfoils during operation and extending the useful life thereof.
The cooling circuits for the stator vanes and rotor blades vary in complexity, performance, and cost of manufacture in a multitude of permutations in a remarkably complex and crowded art due to the decades of development history in modern gas turbine engines.
Even the smallest change in the cooling circuits for these turbine components can have a significant and profound benefit in durability, life, and cost of the resulting turbine component. The various configurations of cooling circuits are due to the different operating environments of the turbine airfoils, including the different velocity, pressure, and temperature distribution over the opposite pressure and suction sides thereof over the span and chord of the airfoils.
Furthermore, a turbine rotor blade is subject to centrifugal forces when rotating with the perimeter of the supporting rotor disk. The turbine blade includes an airfoil extending in radial span from a root at an inner platform to a radially outer tip.
The platforms of the full row of turbine blades define the radially inner flowpath boundary for confining the hot combustion gases at the roots of the airfoils, with an annular turbine shroud confining the combustion gases around the radially outer tips of airfoils.
The platforms are integrally formed at the outer end of a supporting shank which terminates in a supporting dovetail having tangs or lobes configured for being retained in complementary dovetail slots formed in the perimeter of the supporting rotor disk.
The turbine blade is hollow with one or more cooling circuits extending from the base of the dovetail, through the shank and platform and into the airfoil for circulating pressurized air bled from the compressor for cooling the blade during operation.
As indicated above, the cooling circuits inside the turbine airfoil itself may have a myriad of configurations specifically tailored for the operating environment of the turbine airfoil within the combustion gas streamlines. The cooling circuits are fed by inlet channels extending through the shank which receive the pressurized air from the base of the dovetail inside the supporting rotor disk.
The individual blades are retained on the supporting disk typically utilizing forward and aft blade retainers which also provide corresponding seals.
Since the blade shanks and dovetails are shielded from the hot combustion gases by the blade platforms they are readily cooled directly by the cooling air channeled therethrough and by any air surrounding the shanks and dovetails within the sealed cavities of the turbine.
However, the blade platform itself is exposed on its outer side directly to the hot combustion gases and is cooled in part by the cooling air channeled through the internal cooling circuits, as well as in additional part by cooling air channeled under the platform itself. In some configurations, the platform may include conventional film cooling holes extending radially therethrough which bleed a portion of the cooling air from under the platform to its outer surface to create a protective thermally insulating film of cooling air thereover during operation.
The efficiency of a gas turbine engine is increased primarily by increasing the operating temperature of the combustion gases, which correspondingly requires temperature resistant superalloys for the various turbine components and corresponding cooling thereof.
Cooling of the turbine rotor blade platforms is particularly problematic because the platform should be as thin as practical to reduce centrifugal loads carried through the turbine blades, and correspondingly reduce centrifugal stress. Thin platforms render impractical the introduction of internal cooling circuits therein in the myriad of configurations found for the relatively thicker turbine airfoils themselves. And, film cooling of the turbine platforms themselves necessarily has limited effectiveness given the limited amount of cooling air that may be bled from the compressor, which correspondingly reduces overall efficiency of the engine.
Accordingly, it is desired to provide a turbine rotor blade having improved platform cooling for further increasing the efficiency of operation of the gas turbine engine.